The subject matter disclosed herein relates to turbine engines and particularly to methods and apparatus involving shroud cooling in turbine engines.
The high pressure turbine section of a turbine engine includes rotor blades extending radially from a disk assembly mounted inside a casing. The turbine engine includes a shroud assembly mounted on the circumference of the casing surrounding the rotor blades. The rotor blades and shroud assembly are subjected to a high temperature gas flow that affects the rotation of the rotor blades. The rotor blades include a blade tip at a distal end of a rotor blade. A small gap is defined between the blade tips and the shroud assembly. The small gap is desirable for engine efficiency since gas flow passing through the gap does not efficiently affect the rotation of the rotor blades.
In practice, the shroud assembly often comprises a number of segments mounted to the casing to form a circumferential shroud assembly. The shroud assembly is subjected to high temperatures and the segments are often cooled with flowing pressurized air. The pressurized air contacts a surface of a shroud segment and may pass through internal passages of the shroud segment and into the gas flow path inside the casing. Once the pressurized air has cooled the shroud segment, the pressurized air entering the gas flow path may undesirably affect the gas flow path by changing a direction of flow. Thus, it is desirable to reduce the amount of pressurized air used to cool the shroud segment and to discharge the pressurized air into the gas flow path in a manner that lessens the effects to the gas flow path.